Use of auxiliary rudders for yaw control at low speed

ABSTRACT

Apparatus and methods for controlling yaw of a rotorcraft in the event of one or both of low airspeed and engine failure are disclosed. A yaw propulsion provides a yaw moment at low speeds. The yaw propulsion device may be an air jet or a fan. A pneumatic fan may be driven by compressed air released into a channel surrounding an outer portion of the fan. The fan may be driven by hydraulic power. Power for the yaw propulsion device and other system may be provided by a hydraulic pump and/or generator engaging the rotor. Low speed yaw control may be provided by auxiliary rudders positioned within the stream tube of a prop. The auxiliary rudders may one or both of fold down and disengage from rudder controls when not in use.

RELATED APPLICATIONS

This application: is a continuation (divisional) of U.S. patentapplication Ser. No. 13/282,985, filed Oct. 27, 2011 and due to issueFeb. 24, 2015 as U.S. Pat. No. 8,960,594; which claims the benefit ofU.S. Provisional Patent Application Ser. No. 61/409,494, filed Nov. 2,2010; both of which are hereby incorporated by reference in theirentirety.

Additionally, this patent application hereby incorporates by referenceU.S. Pat. No. 5,301,900 issued Apr. 12, 1994 to Groen et al., U.S. Pat.No. 1,947,901 issued Feb. 20, 1934 to J. De la Cierva, and U.S. Pat. No.2,352,342 issued Jun. 27, 1944 to H. F. Pitcairn.

BACKGROUND

The Field of the Invention

This invention relates to rotating wing aircraft, and, more particularlyto rotating wing aircraft relying on autorotation of a rotor to providelift.

The Background Art

Rotating wing aircraft rely on a rotating wing to provide lift. Incontrast, fixed wing aircraft rely on air flow over a fixed wing toprovide lift. Fixed wing aircraft must therefore achieve a minimumground velocity on takeoff before the lift on the wing is sufficient toovercome the weight of the plane. Fixed wing aircraft thereforegenerally require a long runway along which to accelerate to achievethis minimum velocity and takeoff.

In contrast, rotating wing aircraft can take off and land vertically oralong short runways inasmuch as powered rotation of the rotating wingprovides the needed lift. This makes rotating wing aircraft particularlyuseful for landing in urban locations or undeveloped areas without aproper runway.

The most common rotating wing aircraft in use today are helicopters. Ahelicopter typically includes a fuselage, housing an engine andpassenger compartment, and a rotor, driven by the engine, to providelift. Forced rotation of the rotor causes a reactive torque on thefuselage. Accordingly, conventional helicopters require either twocounter rotating rotors or a tail rotor in order to counteract thisreactive torque.

Another type of rotating wing aircraft is the autogyro. An autogyroaircraft derives lift from an unpowered, freely rotating rotor orplurality of rotary blades. The energy to rotate the rotor results froma windmill-like effect of air passing through the underside of therotor. The forward movement of the aircraft comes in response to athrusting engine such as a motor driven propeller mounted fore or aft.

During the developing years of aviation aircraft, autogyro aircraft wereproposed to avoid the problem of aircraft stalling in flight and toreduce the need for runways. The relative airspeed of the rotating wingis independent of the forward airspeed of the autogyro, allowing slowground speed for takeoff and landing, and safety in slow-speed flight.Engines may be tractor-mounted on the front of an autogyro orpusher-mounted on the rear of the autogyro.

Airflow passing the rotary wing, alternately called rotor blades, whichare tilted upward toward the front of the autogyro, act somewhat like awindmill to provide the driving force to rotate the wing, i.e.autorotation of the rotor. The Bernoulli effect of the airflow movingover the rotor surface creates lift.

Various autogyro devices in the past have provided some means to beginrotation of the rotor prior to takeoff, thus further minimizing thetakeoff distance down a runway. One type of autogyro is the “gyrodyne,”which includes a gyrodyne built by Fairey aviation in 1962 and the XV-1convertiplane first flight tested in 1954. The gyrodyne includes athrust source providing thrust in a flight direction and a large rotorfor providing autorotating lift at cruising speeds. To provide initialrotation of the rotor, jet engines were secured to the tip of each bladeof the rotor and powered during takeoff, landing, and hovering.

At high speeds, the direction and orientation of an autogyro may bereadily controlled using conventional control surfaces such as ailerons,rudders, elevators, and the like, that are exposed to air flow over theairframe of the autogyro. Pitch and roll may also be controlled bycyclically altering the pitch of the blades in order to increase thelift at a certain point in the rotation of each blade. Pitch and rollmay also be controlled by altering the angle of the mast coupling therotor to the airframe.

In an emergency landing when an autogyro has lost power, the airspeed ofthe autogyro is likely to be low due to a lack of propulsion. Wherecross winds are present yaw control may be critical in order to maintainthe autogyro aligned with a runway. At low airspeeds, pitch and roll maystill be accomplished using cyclic pitch and mast tilt controls inasmuchas the rotor typically is still auto-rotating.

However, yaw control is not readily accomplished at low air speeds usingconventional control surfaces. Control surfaces, such as rudders, maynot have sufficient airflow thereover at low speeds to induce a yawmoment. In addition, autogyros typically do not have a tail rotorcoupled to the engine to counteract torque exerted by the engine on therotor as do helicopters.

In view of the foregoing, it would be an advancement in the art toprovide means for controlling yaw of an autogyro at low speeds and, inparticular, for controlling yaw of an autogyro in the event of enginefailure.

BRIEF SUMMARY OF THE INVENTION

The invention has been developed in response to the present state of theart and, in particular, in response to the problems and needs in the artthat have not yet been fully solved by currently available apparatus andmethods. The features and advantages of the invention will become morefully apparent from the following description and appended claims, ormay be learned by practice of the invention as set forth hereinafter.

Consistent with the foregoing, a rotorcraft in accordance with anembodiment of the invention may include an airframe, a rotor rotatablymounted to the airframe and rotatable about an axis of rotation, ahorizontal stabilizer mounted to the airframe offset from the axis ofrotation, and an emergency power supply mounted to the airframe. A fanis mounted to the horizontal stabilizer and is selectively coupled tothe emergency power supply to drive the fan in forward and reversedirections in order to generate a yaw moment. In one method of use, theemergency power supply is coupled to the fan in the event that failureof an engine of the rotorcraft is detected.

For example, a method of use may include urging the rotorcrafttranslationally under power of the engine while leaving the fanunpowered. Upon detecting a loss of power of the engine, the emergencypower supply is coupled to the fan in proportion to yaw control inputsfrom a pilot.

In another aspect of the invention, the emergency power supply includesa reservoir of compressed air. In such embodiments, the fan may becoupled to a pneumatic motor for driving the fan. Air from the reservoirmay be supplied to the pneumatic motor in accordance with pilot inputsin order to control yaw of the rotorcraft.

In another aspect of the invention, the fan includes inner blades andouter blades, positioned radially outwardly from the inner blades. Aring extends circumferentially around the inner blades and is positionedradially between the inner blades and outer blades. The outer blades arepositioned within a channel such that the ring and channel form anannular cavity and the ring hinders or restricts escape of air from thechannel. Compressed air from the reservoir may be selectively releasedinto the channel in order to drive the fan.

In another aspect of the invention, a hydraulic motor is coupled to thefan to drive the fan. In such embodiments, the emergency power supplymay include a hydraulic reservoir having a bladder of compressed airtherein for urging the hydraulic fluid out of the reservoir. Fluid fromthe reservoir may be coupled to the hydraulic motor in accordance withpilot inputs in order to drive the fan.

In another aspect of the invention, the emergency power supply includesat least one of a hydraulic pump and a generator selective engageableand rotatably coupled to the rotor. For example, a belt, gear, or otherdrive mechanism may engage the rotor and a drive wheel of one or both ofthe hydraulic pump and generator. A generator may be used to assurepower availability electrical power for any and all aircraft uses ofelectricity, controls, instruments, battery charging, or the like.

A flight control system for performing the above described methods usingthe above described apparatus is also disclosed and claimed herein.

BRIEF DESCRIPTION OF THE DRAWINGS

The foregoing features of the present invention will become more fullyapparent from the following description and appended claims, taken inconjunction with the accompanying drawings. Understanding that thesedrawings depict only typical embodiments of the invention and are,therefore, not to be considered limiting of its scope, the inventionwill be described with additional specificity and detail through use ofthe accompanying drawings in which:

FIG. 1 is an isometric view of an aircraft in accordance with anembodiment of the present invention;

FIG. 2 is a front elevation view of a compressed or otherwisepressurized air supply for a tip jet in accordance with an embodiment ofthe present invention;

FIG. 3A is a front elevation view of a rotor craft illustratingoperational parameters describing a rotor configuration suitable for usein accordance with embodiments of an apparatus and method in accordancewith the present invention and the system of FIGS. 1 and 2 inparticular;

FIG. 3B is a right side elevation view of the rotor craft of FIG. 3A;

FIG. 3C is a partial cut of a right side elevation view of the rotor ofFIG. 3A;

FIG. 4 is an isometric view of an alternative configuration of anaircraft in accordance with an embodiment of the present invention;

FIG. 5A is a partial side elevation view of an aircraft incorporatingair jets for yaw control;

FIG. 5B is a partial top plan cross-sectional view of the aircraft ofFIG. 5A;

FIG. 6A is a partial side elevation view of an aircraft incorporating apneumatic tail fan for yaw control;

FIG. 6B is a partial top plan cross-sectional view of the aircraft ofFIG. 6A;

FIG. 6C is a partial top plan view of a pneumatic tail fan for yawcontrol;

FIG. 7 is a partial side elevation view of an aircraft incorporating amotor-driven tail fan for yaw control;

FIGS. 8A through 8C are schematic block diagrams of emergency powersupplies;

FIGS. 9A and 9B are schematic block diagrams of control systemsconfigured for controlling emergency yaw control propulsion devices;

FIG. 10 is a side elevation view of a mast and hub incorporating anemergency power take-off system;

FIG. 11 is a partial isometric view of drive wheels and hub wheels of anemergency power take-off system;

FIG. 12 is a top plan view of an emergency power take-off system;

FIGS. 13A through 13C are schematic block diagrams of apparatus foractivating an emergency take-off system;

FIG. 14 is a schematic block diagram of a control system for an aircraftincorporating an emergency power take-off system;

FIG. 15 is process flow diagram of a method for operating an aircraftincorporating an emergency yaw control propulsion device;

FIG. 16 is a process flow diagram of a method for operating an aircraftincorporating an emergency power take-off system;

FIG. 17 is a partial isometric view of an aircraft incorporating bothmain and auxiliary rudders;

FIG. 18 is a top plan view illustrating air flow over an aircraftincorporating main and auxiliary rudders;

FIGS. 19A and 19B are isometric views of an auxiliary rudder in deployedand stowed positions, respectively;

FIG. 20 is a schematic block diagram of a control system for an aircraftincorporating both main and auxiliary rudders;

FIG. 21 is a schematic block diagram of an alternative embodiment of acontrol system for an aircraft incorporating both main and auxiliaryrudders;

FIG. 22 is a process flow diagram of a method for operating an aircraftincorporating both main and auxiliary rudders; and

FIG. 23 is a process flow diagram of an alternative method for operatingan aircraft incorporating both main and auxiliary rudders.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS

It will be readily understood that the components of the presentinvention, as generally described and illustrated in the drawingsherein, could be arranged and designed in a wide variety of differentconfigurations. Thus, the following more detailed description of theembodiments of the system and method of the present invention, asrepresented in the drawings, is not intended to limit the scope of theinvention, as claimed, but is merely representative of variousembodiments of the invention. The illustrated embodiments of theinvention will be best understood by reference to the drawings, whereinlike parts are designated by like numerals throughout.

This patent application hereby incorporates by reference U.S. Pat. No.5,301,900 issued Apr. 12, 1994 to Groen et al., U.S. Pat. No. 1,947,901issued Feb. 20, 1934 to J. De la Cierva, and U.S. Pat. No. 2,352,342issued Jun. 27, 1944 to H. F. Pitcairn.

Referring to FIG. 1, an aircraft 10 includes a fuselage 12 defining acabin for carrying an operator, passengers, cargo, or the like. Thefuselage 12 may include one or more fixed wings 14 shaped as airfoilsfor providing lift to the aircraft. The wings 14 may be configured suchthat they provide sufficient lift to overcome the weight of the aircraft10 only at comparatively high speeds inasmuch as the aircraft 10 iscapable of vertical takeoff and landing (VTOL) and does not need liftfrom the fixed wings 14 at low speeds, e.g. below 50 mph or even 100 mphupon taking off.

In this manner, the wings 14 may be made smaller than those of fixedwing aircraft requiring a high velocity takeoff, which results in lowerdrag at higher velocities. In some embodiments the wings 14 providesufficient lift to support at least 50 percent, preferably 90 percent,of the weight of the aircraft 10 at air speeds above 200 mph.

Control surfaces 16 may secure to one or both of the fuselage 12 andwings 14. For example a tail structure 18 may include one or morehorizontal stabilizers 20 and one or more rudders 22. The rudders 22 maybe adjustable as known in the art to control the yaw 24 of the aircraft10 during flight. As known in the art, yaw 24 is defined as rotationabout a vertical axis 26 of the aircraft 10. In the illustratedembodiment, the rudders 22 may comprise hinged portions of thehorizontal stabilizers 20.

The tail structure 18 may further include a horizontal stabilizer 28 andan elevator 30. The elevator 30 may be adjustable as known in the art toalter the pitch 32 of the aircraft 10. As known in the art, pitch 32 isdefined as rotation in a plane containing the vertical axis 26 and alongitudinal axis 34 of the fuselage of an aircraft 10. In theillustrated embodiment, the elevator 30 is a hinged portion of thehorizontal stabilizer 28. In some embodiments, twin rudders 22 may bepositioned at an angle relative to the vertical axis 26 and serve bothto adjust the yaw 24 and pitch 32 of the aircraft 10.

The control surfaces 16 may also include ailerons 36 on the wings 14. Asknown in the art, ailerons 36 are used to control roll 38 of theairplane. As known in the art, roll 38 is defined as rotation about thelongitudinal axis 34 of the aircraft 10.

Lift during vertical takeoff and landing and for augmenting lift of thewings 14 during flight is provided by a rotor 40 comprising a number ofindividual blades 42. The blades are mounted to a rotor hub 44. The hub44 is coupled to a mast 46 which couples the rotor hub 44 to thefuselage 12. The rotor 40 may be selectively powered by one or moreengines 48 housed in the fuselage 12, or adjacent nacelles, and coupledto the rotor 40. In some embodiments, jets 50 located at or near thetips of the blades 42 power the rotor 40 during takeoff, landing,hovering, or when the flight speed of the aircraft is insufficient toprovide sufficient autorotation to develop needed lift.

Referring to FIG. 2, while still referring to FIG. 1, in the illustratedembodiment, the engines 48 may be embodied as jet engines 48 thatprovide thrust during flight of the aircraft. The jet engines 48 mayadditionally supply compressed air to the jets 46 by driving a bypassturbine 62 or auxiliary compressor. Air compressed by the bypass turbine62 may be transmitted through ducts 54 to a plenum 56 in fluidcommunication with the ducts 54.

The plenum 56 is in fluid communication with the mast 46 that is hollowor has another passage to provide for air conduction. A mast fairing 58positioned around the mast 46 may provide one or both of an air channeland a low drag profile for the mast 46. The mast 46 or mast fairing 58is in fluid communication with the rotor hub 44. The rotor hub 44 is influid communication with blade ducts 60 extending longitudinally throughthe blades 42 to feed the tip jets 50.

Referring to FIGS. 3A-3C, rotation of the rotor 40 about its axis ofrotation 72 occurs in a rotor disc 70 that is generally planar but maybe contoured due to flexing of the blades 42 during flight. In general,the rotor disc 70 may be defined as a plane in which the tips of theblades 42 travel. Inasmuch as the blades 42 flap cyclically upward anddownward due to changes in lift while advancing and retreating, therotor disc 70 is angled with respect to the axis of rotation when viewedalong the longitudinal axis 34, as shown in FIG. 3A.

Referring to FIG. 3B, the angle 74 of the rotor disc 70, relative to aflight direction 76 in the plane containing the longitudinal axis 34 andvertical axis 26, is defined as the angle of attack 74 or rotor diskangle of attack 74. For purposes of this application, flight direction76 and air speed refer to the direction and speed, respectively, of thefuselage 12 of the aircraft 10 relative to surrounding air. In autogyrosystems, the angle of attack 74 of the rotor disc 70 is generallypositive in order to achieve autorotation of the rotor 40, which in turngenerates lift.

Referring to FIG. 3C, the surfaces of the blades 42, and particularlythe chord of each blade 42, define a pitch angle 78, or blade angle ofattack 78, relative to the direction of movement 80 of the blades 42. Ingeneral, a higher pitch angle 78 will result in more lift and higherdrag on the blade up to the point where stalling occurs, at which pointlift has declined below a value necessary to sustain flight. The pitchangle 78 of the blade 42 may be controlled by both cyclic and collectivepitch control as known in the art of rotary wing aircraft design.

In the following description power availability is the issue. The powerneed may be mechanical, electrical, pneumatic, hydraulic, or so forth.Regardless, in each of the alternative embodiments, some portioned powerdrawn from the autorotating rotor may be directed to a generatorproviding auxiliary power required to operate the rotorcraft. Likewise,a hydraulic or pneumatic pump by providing alternative, auxiliary powerfor landing gear controls, etc. for “engine-off” flight conditions.

FIG. 4 illustrates an alternative configuration for an aircraft 10. Inthe embodiment of FIG. 4, the airframe 12 includes twin booms 12 a, 12 bextending rearwardly from the fuselage 12. The horizontal stabilizer 28mounts to both of the booms 12 a, 12 b and extends therebetween. Eachboom 12 a, 12 b may also have a vertical stabilizer 20 a, 20 b mountedthereto. The airframe configuration of FIG. 4 may be propelled by jetengines 48 mounted externally as in FIG. 1 or may be propelled by aninternal combustion or turboprop engine 48 mounted internally orexternally to the airframe 12. The engine 48 may rotate a propeller 52or prop 52 in order to propel the aircraft 10 forward. The prop 52 maybe located at the front of the airframe or be located between thehorizontal stabilizer 28 and the front of the airframe 12.

Referring to FIGS. 5A and 5B, in some embodiments, a moment in the yawdirection 24 may be induced by one or more jets 100 a, 100 b offset fromthe center of gravity of the aircraft 10 or the axis of rotation 72 ofthe rotor 40. For example, the jets 100 a, 100 b may be mounted to oneor both of the horizontal stabilizers 20 a, 20 b of FIG. 4 or thevertical stabilizer 20 of FIG. 1. The jets 100 a, 100 b may also bemounted to the fuselage of the airframe 12 offset from the center ofgravity or axis of rotation 72 or to one or both of the booms 12 a, 12b.

The jets 100 a, 100 b may be oriented to direct a jet of air having amajor velocity component directed in a transverse direction 102perpendicular to the longitudinal axis 34 of the aircraft 10. The jet100 a may direct a jet of air in a direction opposite the jet 100 b. Airto the jets 100 a, 100 b may be directed to the jets 100 a, 100 b by oneor more valves 104.

In the illustrated embodiment, a single valve 104 has outputs coupled tothe jets 100 a, 100 b by means of lines 106 a, 106 b, respectively andan input coupled to a source line 108. the valve 104 may be embodied asan electrically or mechanically controlled, three port, two-way,diverter valve operable to control flow of air at variable rates througheither of the lines 106 a, 106 b or neither of the lines 106 a, 106 b,responsive to a control input.

The valve 104 may be embodied as multiple independently controlledvalves configured to have equivalent function to a three port, two-way,diverter valve. The valve 104 directs air from the source line 108responsive to electrical or mechanical control inputs from a pilot orautopilot system in order to create a yaw moment due to reactive forcesat the jets 100 a, 100 b.

Referring to FIGS. 6A and 6B, in an alternative embodiment, one or morefans 110 are mounted to the airframe 12, such as to a verticalstabilizer 20, 20 a, or 20 b. Each fan 110 is positioned within anaperture 112 extending through a portion of the aircraft 10 having anextent parallel to the transverse direction 102, though it may have anextent perpendicular to the transverse direction 102 as well.

In the illustrated embodiment, an aperture 112 extends through thevertical stabilizer 20 a of the aircraft 10 of FIG. 4. An aperture 112may additionally or alternatively extend through the vertical stabilizer20 b. In embodiments such as the aircraft 10 of FIG. 1, an aperture 112may extend through the vertical stabilizer 20. However, an aperture 112may additionally or alternatively extend through the airframe 12 orbooms 12 a, 12 b at some other position offset from the center ofgravity of the aircraft 10 or the axis of rotation 72 of the rotor.

The fan 110 may include inner blades 114 exposed to air within theaperture 112. The fan 110 may additionally include outer blades 116 thatextend internally radially outwardly from the outer diameter of theinner blades 114. The outer blades 116 may project inwardly within theaircraft 10. The inner blades 114 and outer blades 116 may include innerand outer portions of the same blades. In the illustrated embodiment,the outer blades 116 extend internally within the vertical stabilizer20.

A channel 118 may extend around the fan 110 such that the outer blades116 are positioned completely or partially within the channel 118. Aring 120 may extend circumferentially around the inner blades 114positioned between the inner blades 114 and outer blades 116. The ring120 may extend substantially across the channel, e.g. between about 85and 100% of the width of the channel at the location of the ring 120.The ring 120 may serve to substantially hinder leakage of air out of thecircumferential cavity defined by the channel 118 and ring 120.

The lines 106 a, 106 b may be in fluid communication with the channel118 in order to direct air at the outer blades 116 in order to drive thefan 110. The lines 106 a, 106 b may be in fluid communication with ports122 a, 122 b, respectively, in fluid communication with the channel 118.The ports 122 a, 122 b may direct air from the lines 106 a, 106 b suchthat air emitted from the ports 122 a, 122 b emitted into the channel118 has a substantial tangential component with respect to the channel118. The ports 122 a, 122 b may be oriented such that air from a port122 a, 122 b emitted into the channel 118 will have an angular velocitywith respect to the axis of rotation of the fan 110 that is oppositethat of air emitted from the other port 122 b, 122 a.

An outlet port 124 may be in fluid communication with the channel 118 inorder to permit air flow out of the channel 118. Alternatively, thevalve 104 may configured such that the ports 122 a, 82 b may be coupledto ambient air while compressed air is emitted from the other port 82 b,82 a such that an outlet port 124 is not needed.

Referring to FIG. 6C, in some embodiments, the outer blades 116 have achord oriented parallel to the axis of rotation 126 of the fan 110 suchthat air emitted into the channel 118 may more effectively drive theouter blades 116. The inner blades 114 may have a chord oriented at anon perpendicular angle 128 relative to the axis of rotation 126 asknown in the art of prop design in order to more effectively urge airparallel to the axis of rotation 126.

In embodiments where the inner blades 114 and outer blades 116 areportions of the same blade, the blade may be twisted such that a portionof the inner blade 114 has the illustrated angle 128 while a portion ofthe outer blade 116 is parallel to the axis of rotation 126. The chordof the inner blade 114 may have an angle that varies with distance fromthe axis of rotation 126 in order to achieve a desired figure of meritas known in the art of prop design.

Referring to FIG. 7, in an alternative embodiment, the fan 110 iscoupled to and driven by a pneumatic or hydraulic motor 130 mounted tothe airframe 12, such as by mounting the motor 130 to the verticalstabilizer 20. The motor 130 may have inputs 132 a, 132 b coupled tolines 134 a, 134 b coupled to a valve 136. In embodiments where themotor 130 is a hydraulic motor 130, the valve 136 may be additionallycoupled to a source line 138 and a return line 140. The valve 136 may bean electrically or mechanically controlled valve 136 operable to couplethe source line 138 to one of the lines 134 a, 134 b and the return line140 to the other of the lines 134 b, 134 a in response and in proportionto pilot or autopilot inputs.

In embodiments where the motor 130 is a pneumatic motor, a return line140 may be unnecessary and the valve 136 may be a pneumatic valveoperable to couple the source line 138 to one, both, or neither of thelines 134 a, 134 b in response to and in proportion to pilot orautopilot inputs. It may thus drive the fan 110 and create acontrollable yaw moment during one or both of low speed flight andengine failure situations.

Referring to FIG. 8A, pneumatic power for the foregoing yaw controlsdescribed in FIGS. 5A through 7 and corresponding descriptions may besupplied by the illustrated reserve power supply 150. The reserve powersupply 150 may include a compressed air reservoir 152. The reservoir 152may be filled prior to takeoff of the aircraft 10 and be available foruse in the case of an emergency.

Alternatively, a compressor 154 may be coupled to the air reservoir 152.The compressor 154 may receive mechanical or electrical energy derivedfrom the engine 48 or engines 48 in order to maintain pressure withinthe air reservoir 152 above a threshold. The air reservoir 152 may haveother uses during normal operation, e.g., during normal engineoperation, and may additionally serve as an emergency power supply upona loss of engine power.

A valve 156 may couple the air reservoir 152 to the valve 104 or valve136 of the yaw control systems of FIGS. 5 through 7. The valve 156 mayprevent leakage of air from the air reservoir 152 if air were permittedto flow directly from the air reservoir 152 to the valve 104 or valve136.

For example, the valve 104 or valve 136 may be coupled to conventionalyaw controls, e.g., controls for operating a rudder 22, or rudders 22 a,22 b, such that the valve 104 or valve 136 is opened and closedresponsive to these inputs in a way to generate a yaw moment using thejets 100 a, 100 b or fan 110 in at least the same direction as the samecontrol input would induce using the rudder. The valve 156 may ensurethat no air is released from the air reservoir 152 in response to theseinputs until an emergency or other signaling event occurs and the valve156 is opened.

Alternatively, the valve 156 may be omitted and the valves 104 or valve136 may exclusively control flow of air from the air reservoir 152. Insuch an embodiment, pilot controls may be switched in the event of anemergency such that yaw control inputs are coupled to the valve 104 orvalve 136. This coupling may occur instead of, or in addition to, thatof the rudder 22, or rudders 22 a, 22 b. This will provide the yawmoment needed to control the aircraft 10 in the event of one or both oflow airspeeds and engine failure.

Referring to FIG. 8B, in embodiments including a fan 110 powered by ahydraulic motor 130, the reserve power supply 150 may be embodied as ahydraulic reservoir 160. The hydraulic reservoir 160 may contain abladder 162 containing compressed air for driving hydraulic fluid fromthe hydraulic reservoir 160. The bladder 162 may be inflated by acompressor 164 mounted to the airframe 12 and powered by mechanical orelectrical energy derived from the engine 48 or may be filled withcompressed air prior to takeoff of the aircraft 10. The bladder may becontained as a separator in a vessel or may be replaced by a piston.

The volume of the bladder 162 (or equivalent) may be expandable suchthat the compressed air within the bladder 162 tends to urge hydraulicfluid outwardly from the reservoir 120. In some embodiments, a piston164 may be interposed between the bladder 162 and hydraulic fluid withinthe reservoir 160 such that expansion of the bladder 162 urges thepiston against the fluid within the reservoir 160. Alternatively, abladder may act as a separator in a pressure vessel.

As with the embodiment of FIG. 8A, a valve 156 may control flow ofhydraulic fluid to the valve 136 controlling the flow of fluid from thereservoir 162. The valve 156 may be opened in the event of an emergency,as described above with respect to FIG. 8A. Alternatively the valve 136may exclusively control the flow of fluid to the hydraulic motor 130 andyaw control inputs may be coupled to the valve 136 only in the event ofan emergency, as described above with respect to FIG. 8A.

Referring to FIG. 8C, in some embodiments, the air reservoir 152 may beused to power a pneumatic hydraulic pump 166 coupled to a hydraulicreservoir 168. The outlet of the pump 166 may be coupled to the valve136. In such embodiments, the valve 156 may be used to control flow ofair from the reservoir 152 to the pump 166. The valve 156 may be openedin the event of an emergency. In the event of an emergency, the valve156 may be automatically opened and closed in accordance to feedback asto the pressure of fluid downstream of the pump in order to supplysufficient pressure to operate the hydraulic motor 130 responsive topilot inputs to the valve 136.

Referring to FIG. 9A, a control system 180 for an aircraft 10 mayinclude, for example, a control unit 182, pilot controls 184, controlsurface actuators 186, rotor actuators 188, and throttle 190. Thecontrol unit 182 may include conventional avionic computers forperforming navigation and autopiloting functions. The control unit 182may also couple signals from pilot controls 184, such as throttlecontrol, cyclic pitch control, collective pitch control, mast tiltcontrol inputs. The control unit 182 may couple pilot inputs, andcontrols signals generated by an autopilot computer to the controlsurface actuators 186, rotor actuators 188, and the throttle 190.

The control surface actuators 186 may include actuators for actuatingany ailerons 36, elevators 30, rudders 22, 22 a, 22 b, and the like.Rotor actuators 188 may include actuators for controlling mast tilt,cyclic pitch, and collective pitch as known in the art of rotorcraftdesign. In some embodiments, the pilot controls 184 may be coupleddirectly to the control surface actuators 186, rotor actuators 188, andthrottle 190 without an intervening control unit 182.

The control unit 182 may additionally be coupled to the valve 156 of thereserve power supply 150 and the valve 136. The control unit 182 may beoperable to detect a loss of power in the engine 48 and, in response,open the valve 156 and couple pilot inputs relating to yaw control,e.g., rudder controls, to the valve 136, or to the valve 104, in orderto enable to enable the pilot to control yaw of the aircraft 10. Thecontrol unit 182 may be coupled to sensors 192 within the engine 48 orto structures driven by the engine 48 in order to detect whether theengine 48 is outputting sufficient power, or otherwise as known in theart of engine and aircraft design.

In some embodiments, the control unit 182 may open the valve 156 andcouple pilot inputs relating to yaw control to the valve 136, or thevalve 104, only upon detecting an airspeed below a threshold, or onlyupon detecting an airspeed below a threshold and a loss of power in theengine. The velocity may be determined by means of Global PositioningSystem (GPS) data, by means of an airspeed sensor coupled to the controlunit by both, or the like.

In still other embodiments, opening of the valve 156 may additionally oralternatively be performed by the pilot either directly or by providingan input to the control unit 182, which then actuates the valve 156.Coupling of yaw control inputs to the valve 104 or valve 136 may also beperformed by means of a manual operation of a switch, valve, or someother actuator. A pilot may decide to engage the emergency operationrather than rely on automatic actuation.

Referring to FIG. 9B, in some embodiments, the control unit 182selectively switches yaw control signals from the rudder 22 or rudders22 a, 22 b to the valve 136 or 104. The yaw control signals may beinitiated by the pilot controls 184 or by an autopilot function of thecontrol unit 182. The control unit 182 may be programmed to detect acondition, such as loss of air speed or power in the engine 48 and, inresponse, couple yaw control signals to the valve 136 or the valve 104.

In some embodiments, the control unit 182 may couple yaw control signalsto the valve 136 or valve 104 only upon detecting an airspeed below athreshold, or only upon detecting both an airspeed below a thresholdcombined with a loss of power in the engine. The velocity may bedetermined by means of Global Positioning System (GPS) data or by meansof an airspeed sensor coupled to the control unit. In some embodiments,the switching of yaw controls to the valve 136 or valve 104 may beperformed manually by means of a switch included among the pilotcontrols 184. Thus sophisticated controls may be replaced by pilotjudgment and manual controls.

Referring to FIG. 10, in an autogyro, the rotor 40 is typically orientedsuch that the rotor 40 is compelled to rotate by lift and drag forces onthe blades 42 in response to air flow over the rotor 40 due to forwardmovement of the aircraft. Therefore, in the event of a loss of power ofthe engine 48, the forward momentum of the aircraft 10 may cause therotor 40 to continue to rotate. Accordingly, in some embodiments, anemergency power take-off 200 is coupled to the rotor 40 in order toprovide one or more of hydraulic, pneumatic, and electrical power forthe systems of the aircraft 10. The emergency power take-off 200 mayprovide one or both of hydraulic and pneumatic power to the valve 104 orvalve 136 in order to provide yaw control in the event of a loss ofpower and at low speeds.

In the illustrated embodiment, a shroud 202 surrounds the mast 46 and isin fluid communication with the plenum 56. The space between the mast 46and shroud 202 is in fluid communication with a cavity 206 defined bythe hub 44 and in fluid communication with the blade ducts 60. The powertake-off 200 is mounted to the shroud, such as by means of securement toa flange 208 mounted to the shroud 202. Other mounting configurationsare also possible. For example, the power takeoff 200 may mount directlyto the mast 46 or to a flange secured to the mast 46.

Referring to FIGS. 11 and 12, the power take-off 200 may include one ormore drives, such as drive wheels 210 that engage a hub wheel 212. Thehub wheel 212 may be embodied as a portion of the hub 44 or a ringsecured to the hub 44. The drive wheel 210 may engage the hub wheel 212directly or by means of a belt. In the illustrated embodiment, the drivewheel and hub wheel 212 are grooved to reduce slippage of the belt withrespect to the wheels 210, 212.

Where the drive wheel 210 engages the hub wheel 212 directly, the drivewheel 210 and hub wheel 212 may each include a geared surface or aresilient ring creating a high friction interface between the wheels210, 212. In the illustrated embodiment, the drive wheel 210 engages anouter surface of the hub wheel 212. However, in some embodiments, thedrive wheel 210 may engage an inner surface of the hub wheel 212 and bepositioned within the hub wheel 212.

In the illustrated embodiment, the power take-off 200 includes two drivewheels 210, each coupled to one of a generator 214 and a hydraulic pump216. In some embodiments, a pneumatic compressor may be used in theplace of one or both of the generator 214 and the hydraulic pump 216.Alternatively, electrical or hydraulic power from the generator 214 orhydraulic pump 216 may drive a compressor 154 for providing air to avalve 104 or valve 136.

The generator 214 and hydraulic pump 216 may be mounted to the flange208 opposite the drive wheels 210 such that drive shafts of thegenerator 214 and hydraulic pump 216 extend through the flange 208 tocouple the generator 214 and hydraulic pump 216 to the drive wheels 210.Referring specifically to FIG. 12, a tension arm 220 may also bepivotally secured to the flange 208 and be biased, such as by means of aspring, to urge a roller 222 against a belt 218 encircling the drivewheels 210 and hub wheel 212.

Referring to FIGS. 13A through 13C, various means may be used toactivate and deactivate the power take-off 200. In order to avoidunnecessary loss of power during normal flying conditions, it may beadvantageous to deactivate the power take-off 200. Referringspecifically to FIG. 13A, in some embodiments a mechanically,electrically, or hydraulically actuated clutch 230 is interposed betweena drive wheel 210 and the generator 214. A clutch 230 may be similarlyinterposed between a drive wheel 210 and the hydraulic pump 214 or apneumatic compressor.

Referring to FIG. 13B, alternatively, the generator 214 may be activatedand deactivated by controlling the amount of current applied to fieldcoils 232 within the generator 214. As known in the art of generatordesign, electric current is generated by moving a coil through amagnetic field. The magnetic field may be generated by permanentmagnets, electromagnets, or a combination of the two. Accordingly, thedrag induced by the generator 214 may be reduced when not in use byturning off current to field coils 232 used as electromagnets forgenerating a magnetic field. During an emergency situation, current fromthe generator 214 may be routed to the field coils 232 in order toincrease the current output from the generator 214.

Referring to FIG. 13C, similarly, when the hydraulic pump 216 is not inuse, a valve 234 may prevent the flow of hydraulic fluid into thehydraulic pump 216 or route fluid through a recirculation path 236 inorder to reduce drag induced by the hydraulic pump 216 when not in use.When in use, fluid may be routed in and out of the hydraulic pump 216 bymeans of input and output lines 238 a, 238 b, respectively.

Referring to FIG. 14, the control unit 182 may be configured to handleactivation of the emergency power take-off 200 in the event of a loss ofpower due to engine failure or failure of one or both of a primarygenerator 242 or primary hydraulic pump 244. The control unit 182 may beprogrammed to activate a clutch 230 coupling a drive wheel 210 to thegenerator 214 or hydraulic pump 216 in response to detection of a lossof power from the engine 48 or a loss of voltage or pressure from aprimary generator 242 or primary hydraulic pump 244, respectively.

Alternatively, the control unit 182 may be configured to supply power tofield coils 232 or redirect fluid away from a recirculation path 236.This may occur in response to a detection of a loss of power from theengine 48 or a loss of voltage or pressure from a primary generator 242or primary hydraulic pump 244, respectively.

The control unit 182 may be further programmed to activate switches andvalves necessary to route current or pressurized hydraulic fluid to thesystems of the aircraft 10, such as the control surface actuators 186and rotor actuators 188, from the generator 214 or hydraulic pump 216 asneeded. This may include routing hydraulic fluid to a motor 130 poweringthe fan 110 or to a pneumatic pump for driving the fan 110 or jets 100a, 100 b in the embodiments of FIGS. 3 through 5. The control unit 182may further be programmed to sequester the primary generator 242 orprimary hydraulic pump 244 to avoid leakage of current or fluid or toreduce power loss. Some or all of the operations described above asbeing performed by the control unit 182 with respect to the powertake-off 200 may also be performed by a pilot manually operatingswitches, valves, or other mechanical actuators.

Referring to FIG. 15, the foregoing apparatus may be used to perform theillustrated method 250 to achieve yaw control in the event of one orboth of loss of engine power and low speed flight. The method 250 may beperformed partially or completely by means of one or both of a controlunit 182 or pilot inputs. The method 250 includes detecting 252 loss ofyaw control. This may include detecting 252 the occurrence of one or allof, either separately or simultaneously, loss of power from the engine48, loss of control of a rudder 22, and an air speed below a thresholdair speed.

In the event of detection 252 of a loss of yaw control, a power supplyis coupled 254 to a yaw propulsion device. The yaw propulsion device mayinclude jets 100 a, 100 b, a reversible tail fan 110, or the like. Thepower supply may be an emergency power supply such as one of thosedescribed above with respect to FIGS. 8A through 8C. In suchembodiments, coupling 254 the power supply to the yaw propulsion devicemay include opening the valve 156 as described hereinabove.

The power supply may also be an emergency power take-off 200 asdescribed above with respect to FIGS. 10 through 14. In suchembodiments, coupling 254 the power supply to the yaw propulsion devicemay include activating one or both of a secondary generator 214 orsecondary hydraulic pump 216 as described hereinabove. In someembodiments, the power supply may be a pneumatic compressor coupled tothe rotor 40 in the same manner as the hydraulic pump 216. Accordingly,coupling 254 the power supply to the yaw propulsion device may includecoupling the compressor to one of the valve 104 and the valve 136.

In the event that a loss of yaw control is due to low air speed, thepower supply systems coupled 254 to the yaw propulsion device may be theaircraft's primary power systems, such as a primary hydraulic pump 244,electric generator 242, or a pneumatic compressor, powered by power fromthe engine 48. Pilot inputs, or autopilot control signals, may then becoupled 256 to the yaw propulsion device, such as by coupling the yawcontrol signals to a valve 104 or valve 136. The pilot inputs maycontinue to be coupled to the rudder 22 as well or may be completelyredirected to the yaw propulsion device.

Referring to FIG. 16, the illustrated method 260 may be executed inresponse to a loss of power due to failure of an engine 48 or failure ofa primary generator 244 or hydraulic pump 244. The method 260 may beexecuted wholly or completely by one or both of the control unit 182 andpilot inputs. The method 260 includes detecting 262 a loss of power ofan engine 48, primary generator 242, or primary hydraulic pump 244.

If engine failure or primary generator failure is detected 262, then thesecondary generator 214 is activated 264, such as by engaging a clutch230 or supplying current to field coils 232. If engine failure orprimary hydraulic pump failure is detected, then the secondary hydraulicpump 216 is activated 266, such as by engaging a clutch 230 or routinghydraulic fluid away from a recirculation path 236 and through input andoutput lines 238 a, 238 b.

Power from one or both of the secondary generator 214 and secondaryhydraulic pump 216 may then be routed through the systems of theaircraft 10, including the yaw propulsion systems described in FIGS. 5Athrough 9B. Other systems may include devices or actuators, such as thecontrol surface actuators 186, rotor actuators 188, landing gearactuators, electric servos, batteries, instruments, and the like.

Referring to FIG. 17, in some embodiments, an aircraft 10, such as theaircraft 10 of FIG. 4, may include auxiliary rudders 280 mounted to thehorizontal stabilizer 28 and positioned between the rudders 22. In theillustrated embodiment, rudders 280 are mounted to opposing surfaces ofthe horizontal stabilizer 28 (e.g., horizontal airfoil 28 stabilizingvertically as an elevator and stabilizer do in a fixed wing aircraft).They project upward and downward from the horizontal stabilizer 28 inthe vertical direction 26.

Referring to FIG. 18, in flight, the rudders 22 are exposed to air flow282 which is in the “free stream” or substantially undisturbed ambientair through which the aircraft 10 passes. Therefore, the flow 282 has avelocity relative to the aircraft 10. Actuation of the rudders 22 causesthe rudders 22 to redirect a portion of this air flow 282, resulting inmomentum change and a force, causing a yaw moment on the aircraft 10. Atlow speeds, especially in the event of an “engine off” landing, the airflow 282 may be too slow to provide an adequate yaw moment, particularlyin the presence of cross winds.

The auxiliary rudders 280 advantageously are located within a propstream tube 284, which includes air flow impelled by the prop 284, ajet, or some other propulsion source. Inasmuch as the velocity of airwithin the prop stream tube 284 is independent of the airspeed of theaircraft 10, the velocity of airflow over the auxiliary rudders 280 maybe larger than the velocity of air incident on the rudders 22. Theauxiliary rudders 280 therefore may be able to generate a yaw momentgreater than that generated by the rudders 22 at a given air speed.

Referring to FIGS. 19A and 19B, while still referring to FIG. 17,inasmuch as the auxiliary rudders 280 are not needed during high speedflight, the rudders 280 may advantageously fold to a lower profileconfiguration during high speed flight, as shown by the dottedrepresentation of the rudders 280 in FIG. 17. Various methods of hingingand actuation may be used to fold the rudders 280 during high speedflight as known in the art of mechanical design.

Hydraulic, electrical, and mechanical actuators as known in the art ofaircraft design for actuating control surfaces may be used. For example,a deployment shaft 290 may be rotatably mounted within the horizontalstabilizer 28 and be actuated by means of an actuator such as by meansof hydraulic drives or one or more deployment cables 292 a, 292 b thatare tensioned and relaxed to alter the orientation of an auxiliaryrudder 280.

A lock 294 may retain the rudder 280 in the deployed position. Anysuitable locking mechanism known in the mechanical art may be used. Forexample, an actuated piston 296 within a cylinder 298 may be driven bymeans of electrical, hydraulic, or pneumatic power into a receptacle 300formed in the deployment shaft 290 in order to retain the rudder 280 inthe deployed position. The rudder 280 may be rotatably mounted to thedeployment shaft 290 such that following deployment the rudder 280 maybe rotated in order to induce a yaw moment on the aircraft 10.

In the illustrated embodiment, the rudder 280 rotatably mounts to ashaft 302 extending perpendicular to the deployment shaft 290. Therudder 280 may rotatably mount to the shaft 302 or the shaft 302 mayrotatably mount to the deployment shaft 290. An actuator, such as cables306 a, 306 b may engage the shaft 302 or rudder 280 and may be actuatedin order to change the angle of the rudder 280 within the prop streamtube 288.

Referring specifically to FIG. 19B, the deployment shaft 290 may beactuated, such as by means of the cables 292 a, 292 b, or some otheractuator, in order to move the rudder 280 into the illustrated stowedposition in which the rudder 280 is oriented substantially parallel,e.g. within about 10 degrees, to the horizontal stabilizer 28.

For example, the axis of rotation of the rudder 280 about the shaft 302may be more parallel to the horizontal stabilizer 28 than when therudder 280 is in the deployed position. The angular separation betweenthe deployed and stowed positions may be between about 70 and 100degrees.

The horizontal stabilizer 28 may define a receptacle 308 or recess 308for receiving all or part of the rudder 280. Thus, an exposed surface310 of the rudder 280 projects less prominently from the horizontalstabilizer 28. The lock 294, or some other lock, may also retain therudder 280 in the stowed position. For example, the piston 296 may beurged into a stowage receptacle 312 formed in the deployment shaft 250in order to retain the rudder 280 in the stowed position.

Referring to FIG. 20, in some embodiments, the control unit 182 iscoupled to a main rudder actuator 320 operable to change the orientationof the rudders 22 a, 22 b. The control unit may also be coupled to anauxiliary rudder actuator 322. In some embodiments, the auxiliaryrudders 280 are actuated synchronously with the rudders 22 a, 22 b atall times. In such embodiments, the rudders 22 a, 22 b and auxiliaryrudders 280 may be actuated by common actuators and linked to oneanother such that they are compelled to move in unison.

Referring to FIG. 21, in an alternative embodiment, a switch 324controls which of the rudders 22 a, 22 b and the auxiliary rudders 280receives yaw control inputs. The switch 324 may couple control signalsto one or both of the main rudder actuator 320 and auxiliary rudderactuator 322 depending on the state of the switch 324. Alternatively,the switch 324 may change the coupling of mechanical, pneumatic, orhydraulic force from a single actuator to one or both of the rudders 22a, 22 b and the rudders 280 according to the state of the switch 324.

The control unit 182 may include a rudder selector 326 programmed tooperate the switch 324. The rudder selector 326 may be programmed tocouple yaw control inputs from one or both of the pilot controls 184 andan autopilot computer to the auxiliary rudders 280 when the airspeed ofthe aircraft 10 is below a threshold and to couple yaw control inputs tothe rudders 22 a, 22 b when the airspeed of the aircraft 10 is above thethreshold.

In one embodiment a transition region is defined such that both therudders 22 a, 22 b and rudders 280 are actuated simultaneously forairspeeds within the transition region. The rudders 280 are actuatedexclusively below the transition region. The rudders 22 a, 22 b areactuated exclusively above the transition region. In some embodiments,the pilot inputs 184 may additionally or alternatively include manuallyoperable interface to control the switch 324 and select one or both ofthe rudders 22 a, 22 b and rudders 280 to receive yaw control inputs.

In some embodiments, an auxiliary rudder extender 328 may actuate therudders 280 to transition the rudders 280 between the stowed anddeployed orientations described hereinabove. In such embodiments, thecontrol unit 182 may include an auxiliary rudder deployment controller330.

The deployment controller 330 may be programmed to move the auxiliaryrudders 280 to the deployed orientation when the controller 182determines that yaw control inputs are to be coupled to the auxiliaryrudders 280 as described hereinabove. The deployment controller 330 mayalso be programmed to move the auxiliary rudders 280 to the stowedorientation when the controller 182 determines the yaw control inputsare too coupled to the rudders 22 a, 22 b as described hereinabove. Thepilot inputs 184 may also include an interface to control the auxiliaryrudder extender 328 in addition or as an alternative to the auxiliaryrudder deployment controller 330 of the control unit 182.

Referring to FIG. 22, a pilot, control unit 182, or a combination of thetwo, may execute a method 340 for operating an aircraft 10 includingboth main rudders 22 a, 22 b and one or more auxiliary rudders 280. Themethod 340 may include evaluating 342 the airspeed of the aircraft 10with respect to a threshold. If the air speed is below the threshold,then the auxiliary rudders 280 are deployed 344, such as by moving therudders 280 to the deployed position.

In embodiments where the rudders 280 are not movable between stowed anddeployed positions, deployment 344 may be omitted. Yaw control inputsfrom a pilot, or an autopilot computer, may then be coupled to theauxiliary rudders 280 either synchronously with the main rudders 22 a,22 b, exclusive of the main rudders 22 a, 22 b, or in some other controlscheme optimizing used each.

If the airspeed is above the threshold, then the rudders 280 may bestowed 348. This is useful in embodiments having rudders 280 movablebetween deployed and stowed positions. Yaw control inputs from a pilotor autopilot computer may then be coupled 350 to the main rudders 22 a,22 b and decoupled from the auxiliary rudders 280.

Referring to FIG. 23, in an alternative embodiment, a pilot, controlunit 142, or a combination of the two, may execute a method 360 foroperating an aircraft 10 including both main rudders 22 a, 22 b and oneor more auxiliary rudders 280. The method 360 includes evaluating 362whether the airspeed of the aircraft 10 is above a transition region. Ifso, then the auxiliary rudders 280 are stowed 364 if they are found in adeployed position and are movable between deployed and stowed positions.Rudder control inputs are then coupled 366 exclusively to the mainrudders 22 a, 22 b.

The method 360 may further include evaluating 368 whether the airspeedof the aircraft 10 is within the transition region. If so, then theauxiliary rudders 280 are deployed 370 if they are found in the stowedposition and if the rudders 280 are movable between stowed and deployedpositions. Rudder control inputs are then coupled 372 to both theauxiliary rudders 280 and main rudders 22 a, 22 b.

The method 320 may further include evaluating 374 whether the airspeedof the aircraft 10 is below the transition region. If so, then theauxiliary rudders 280 are deployed 376 if they are not already in thedeployed position and if they are movable between stowed and deployedpositions. Rudder control inputs are then coupled 378 to the auxiliaryrudders either exclusive of or synchronously with the main rudders 22 a,22 b.

The present invention may be embodied in other specific forms withoutdeparting from its spirit or essential characteristics. The describedembodiments are to be considered in all respects only as illustrative,and not restrictive. The scope of the invention is, therefore, indicatedby the appended claims, rather than by the foregoing description. Allchanges which come within the meaning and range of equivalency of theclaims are to be embraced within their scope.

What is claimed and desired to be secured by United States LettersPatent is:
 1. A rotorcraft comprising: an airframe including a cabinportion supporting flight controls and a seat to receive a pilot, atower portion fixed to the cabin portion and extending upward therefrom,an empennage extending from the cabin portion away therefrom, a firstvertical stabilizer mounted to the empennage, a first rudder secured topivot with respect to the first vertical stabilizer, and a horizontalstabilizer connected to the empennage; a rotor rotatably mounted to thetower to rotate about an axis of rotation offset from the first verticalstabilizer and the horizontal stabilizer; a propulsion source mounted tothe airframe and directing air flow across the vertical stabilizer; afirst auxiliary rudder mounted to the horizontal stabilizer and spacedfrom the first vertical stabilizer; and a controller receiving yawcontrol inputs from a pilot and selectively actuating at least one ofthe first rudder and the first auxiliary rudder in response to the yawcontrol inputs; wherein the first auxiliary rudder is positioned to bewithin a stream tube of the propulsion source during flight and thefirst rudder is not positioned to be within to be the stream tube of thepropulsion source during flight.
 2. The rotorcraft of claim 1, whereinthe first auxiliary rudder is hinged to the horizontal stabilizer andpivotable with respect thereto between a deployed position, extendingsubstantially vertically, and a stowed position, substantially alignedwith and aerodynamically proximate the horizontal stabilizer.
 3. Therotorcraft of claim 2, wherein the controller is computerized andprogrammed to control movement of the first auxiliary rudder from thestowed position to the deployed position upon detecting an airspeed ofthe aircraft below the airspeed threshold.
 4. The rotorcraft of claim 2,wherein the vertical stabilizer further comprises a receptacle sized toreceive at least a portion of the first auxiliary rudder when the firstauxiliary rudder is in the stowed position.
 5. The rotorcraft of claim1, wherein the first auxiliary rudder is one of a plurality of auxiliaryrudders offset from one another and on opposite surfaces of thehorizontal stabilizer.
 6. The rotorcraft of claim 1, further comprising:a second empennage extending away from the cabin portion; a secondvertical stabilizer secured to the second empennage; and a second rudderpivotably mounted to move with respect to the second verticalstabilizer.
 7. The rotorcraft of claim 1, further comprising: thepropulsion source centrally mounted to drive an air stream away from thecabin portion and toward the horizontal stabilizer, the first and secondrudders spaced apart and pivoting with respect to the first and secondvertical stabilizers, respectively.
 8. The rotorcraft of claim 7,further comprising: a second auxiliary rudder, spaced from the firstauxiliary rudder, the first and second auxiliary rudders operableindependently from the first and second rudders.
 9. The rotorcraft ofclaim 8, further comprising: cooperative linkage connecting the firstand second auxiliary rudders to move in coordination with one another;the first and second auxiliary rudders positioned to re-direct a centralportion of the airstream therebetween upon pivoting of the first andsecond auxiliary rudders; and the first and second auxiliary rudderspositioned to fit at least partially into recesses fitted thereto andformed in the vertical stabilizer.
 10. A rotorcraft comprising: anairframe including first and second vertical stabilizers mounted to theairframe, each of the first and second horizontal stabilizers having arudder mounted thereto, and a horizontal stabilizer mounted to theairframe between the first and second vertical stabilizers; a rotorrotatably mounted to the airframe and rotatable about an axis ofrotation, the axis of rotation offset from the horizontal stabilizer andfirst and second vertical stabilizers; a propulsion system mounted tothe airframe and oriented to direct air flow over the verticalstabilizer, the propulsion system defining a stream tube; an auxiliaryrudder mounted to the horizontal stabilizer positioned between the firstand second vertical stabilizers; and a controller programmed to receiveyaw control inputs from a pilot and to actuate the auxiliary rudder inresponse to the yaw control inputs; wherein the auxiliary rudder ispositioned within the stream tube of the propulsion system and therudder is not positioned within the stream tube of the propulsionsystem.
 11. The rotorcraft of claim 10, wherein the auxiliary rudder ishingedly mounted to the horizontal stabilizer and is rotatable between astowed position and a deployed position, the auxiliary rudders beingoriented parallel to the horizontal stabilizer in the stowed position.12. The rotorcraft of claim 11, wherein the controller is programmed toacutate the auxiliary rudder from the stowed to the deployed positionupon detecting an airspeed of the aircraft below the airspeed threshold.13. The rotorcraft of claim 11, wherein the horizontal stabilizerfurther comprises a receptacle sized to receive at least a portion ofthe auxiliary rudder when the auxiliary ruder is in the stowed position.14. The rotorcraft of claim 10, wherein the auxiliary rudder is one of aplurality of auxiliary rudders positioned offset from one another and onopposing surfaces of the horizontal stabilizer.
 15. A rotorcraftcomprising: an airframe including a cabin portion supporting flightcontrols and a seat to receive a pilot, a tower portion fixed to thecabin portion and extending upward therefrom, first and secondempennages extending from proximate the cabin portion away therefrom,first and second vertical stabilizers mounted to the first and secondempennages, respectively, first and second rudders secured to pivot withrespect to the first and second vertical stabilizers, respectively, anda horizontal stabilizer connected to extend between the first and secondempennages; a rotor, rotatably mounted to the tower to rotate about anaxis of rotation offset from the horizontal stabilizer and the first andsecond vertical stabilizers; a propulsion source mounted to the airframeand directing air flow across the horizontal stabilizer; a firstauxiliary rudder mounted to the horizontal stabilizer and spaced fromthe first and second vertical stabilizers; and a controller receivingyaw control inputs from the pilot and selectively actuating the firstauxiliary rudder in response to the yaw control inputs; wherein— thepropulsion source defines a stream tube extending rearwardly therefrom;the first auxiliary rudder is positioned within the stream tube; and thefirst rudder and first vertical stabilizer are not positioned within thestream tube.
 16. The rotorcraft of claim 15, wherein the first auxiliaryrudder is hinged to the horizontal stabilizer and pivotable with respectthereto between a deployed position, extending substantially vertically,and a stowed position, substantially aligned with and aerodynamicallyproximate the horizontal stabilizer.
 17. The rotorcraft of claim 15,wherein: the controller is computerized and programmed to controlmovement of the first auxiliary rudder from the stowed position to thedeployed position upon detecting an airspeed of the aircraft below theairspeed threshold; the horizontal stabilizer further comprises areceptacle sized to receive at least a portion of the first auxiliaryrudder when the first auxiliary rudder is in the stowed position; thehorizontal stabilizer is secured to extend between the first and secondempennages; the propulsion source is centrally mounted to drive an airstream away from the cabin portion and toward the horizontal stabilizer;a second auxiliary rudder is spaced from the first auxiliary rudder andthe first and second auxiliary rudders are connected to be operableindependently from the first and second rudders, and positioned tore-direct a central portion of the airstream therebetween upon pivotingof the first and second auxiliary rudders; and the first and secondauxiliary rudders are positioned to fit at least partially into recessesfitted thereto and formed in the horizontal stabilizer.
 18. Therotorcraft of claim 15, further comprising a sensor, connected to senseairspeed of the rotorcraft and provide to the controller a parameterreflecting the airspeed.
 19. The rotorcraft of claim 15, wherein thehorizontal stabilizer further comprises a recess formed to receive thefirst auxiliary rudder at airspeeds above a threshold airspeed.